Method for forming a coating matrix on a shaft and disk assembly for a turbine

ABSTRACT

A method for forming a coating matrix on a bore surface of a turbine disk wherein the coating matrix is applied at an interface between the disk and a turbine shaft. The coating matrix enhances thermal conductivity to increase heat transfer from the disk. The method includes providing a receiving surface on the bore surface. The receiving surface is then heated to melt the receiving surface. Next, at least one coating matrix layer is deposited on the receiving surface. The coating matrix layer includes a graphene layer. A pulsed laser system or a robot welding system may be used to melt the receiving section.

FIELD OF THE INVENTION

This invention relates to shrink fit arrangements for turbine components, and more particularly, to a method for forming a thermal conductivity coating matrix on a component of the shrink fit arrangement.

BACKGROUND OF THE INVENTION

Renewable energies and their inherent intermittent character require faster start-up times for conventional power generation methods, which use gas or steam turbines. However, conventional shrink fit arrangements used to fasten or attach components in such turbines limit start-up speed. Referring to FIG. 1, a partial cross sectional view of an exemplary compressor section 10 of a gas turbine is shown. Air enters an inlet 12 and is guided to the compressor section 10 by inlet guide vanes 14. The compressor section 10 includes one or more disks 16 each including a plurality of outwardly extending compressor blades 18. The disks 16 are each attached to a rotor shaft 20 of the gas turbine.

Referring to FIG. 2, a partial cross sectional view of an exemplary shaft 20 and disc 16 is shown. The shaft 20 extends through a central bore 22 formed in the disk 16. The bore 22 and the shaft 20 are dimensioned to form a shrink fit arrangement. The disk 16 is assembled to the shaft 20 by first heating the disk 16, which increases the size of the bore 22. The shaft 20 is then inserted through the bore 22 in the disk 16. The disk 16 is then cooled thus capturing the shaft 20 and forming a shrink fit interface 24 between the bore 22 and the shaft 20. The shrink fit generates stresses on the shaft 20, which serve to attach the disk 16 to the shaft 20. Shrink fit arrangements are also used to attach disks to a rotor shaft in a steam turbine. An outer diameter 21 of the disk 16 includes a plurality of channels or blade grooves 23 (only three grooves 23 are shown in FIG. 2 for purposes of clarity). Each groove 23 is adapted to receive a mating projecting portion of an associated compressor blade 18 thus attaching the compressor blade 18 to the disk 16.

Conventional start-up of a gas turbine results in heating of the disk 16, which decreases shrink fit stress. The shrink fit arrangement is designed such that an amount of shrink fit stress remains, even after the disk 16 has been heated, which is sufficient for maintaining attachment of the disk 16 to the shaft 20. However, a faster start-up speed increases the rate at which the disk 16 is heated. When this occurs, heat cannot be transferred from the disk 16 at a sufficient rate to provide adequate shrink fit stress to keep the disk 16 attached to the shaft 20, resulting in possible turbine failure. Alternatively, the shrink fit arrangement between the disks 16 and the shaft 20 could be changed to a bolted design. However, this increases costs and results in an extended down time in order to implement the bolted design for turbines, which are already in use.

SUMMARY OF INVENTION

A method is disclosed for forming a coating matrix on a first component used in a shrink fit arrangement wherein the coating matrix is applied at an interface between the first component and a second component. The coating matrix enhances thermal conductivity to increase heat transfer from the first component. The method includes providing a receiving section on the first component. The receiving section is then heated in order to melt the receiving section. Next, at least one coating matrix layer is deposited on the receiving section. The coating matrix layer includes a graphene layer. A pulsed laser system or a robot welding system may be used to melt the receiving section. The method may be used to form a coating matrix on a bore surface of a turbine disk or on a turbine shaft. In addition, the coating matrix may also be formed on first and second sides of the disk, a disk outer diameter and on grooves formed on the outer diameter used to attach associated blades.

The respective features of embodiments of the present invention may be applied jointly or severally in any combination or sub-combination by those skilled in the art.

BRIEF DESCRIPTION OF DRAWINGS

The teachings of the present invention can be readily understood by considering the following detailed description in conjunction with the accompanying drawings, in which:

FIG. 1 is a partial cross sectional view of an exemplary compressor section of a gas turbine.

FIG. 2 is a partial cross sectional view of an exemplary shaft and disc of a turbine.

FIG. 3 is schematic depiction of a gas turbine.

FIG. 4 is a partial cross sectional view of a shaft and disc of a turbine and depicts a coating matrix located at a shrink fit interface in accordance with the invention.

FIG. 5 is a cross sectional view of the shaft and disk of a turbine and depicts the coating matrix located on first and second side surfaces, the outer diameter and a groove of the disk.

FIG. 6 is a cross sectional view of the disk without the shaft and depicts a receiving section.

FIG. 7 depicts a first coating matrix formed on the receiving section.

FIG. 8 depicts second and third coating matrixes formed on the receiving section.

FIGS. 9-10 depict alternate embodiments of the coating matrix.

To facilitate understanding, identical reference numerals have been used, where possible, to designate identical elements that are common to the figures.

DETAILED DESCRIPTION

Although various embodiments that incorporate the teachings of the present invention have been shown and described in detail herein, those skilled in the art can readily devise many other varied embodiments that still incorporate these teachings. The invention is not limited in its application to the exemplary embodiment details of construction and the arrangement of components set forth in the description or illustrated in the drawings. The invention is capable of other embodiments and of being practiced or of being carried out in various ways. Also, it is to be understood that the phraseology and terminology used herein is for the purpose of description and should not be regarded as limiting. The use of “including,” “comprising,” or “having” and variations thereof herein is meant to encompass the items listed thereafter and equivalents thereof as well as additional items. Unless specified or limited otherwise, the terms “mounted,” “connected,” “supported,” and “coupled” and variations thereof are used broadly and encompass direct and indirect mountings, connections, supports, and couplings. Further, “connected” and “coupled” are not restricted to physical or mechanical connections or couplings.

Referring to FIG. 3, a gas turbine 26 is schematically shown. The turbine 26 includes a compressor 28, which draws in ambient air 30 and delivers compressed air 32 to a combustor 34. A fuel supply 36 delivers fuel 38 to the combustor 34 where it is combined with the compressed air 32 and the fuel 38 is burned to produce high temperature combustion gas 40. The combustion gas 40 is expanded through a turbine section 42, which includes a series of rows of stationary vanes and rotating turbine blades. The combustion gas 40 causes the blades to rotate to produce shaft horsepower for driving the compressor 28 and a load, such as an electrical generator 44. Expanded gas 46 is either exhausted to the atmosphere directly, or in a combined cycle plant, may be exhausted to atmosphere through a heat recovery steam generator.

Referring to FIG. 4, it is desirable to transfer heat from the previously described disk 16 at a sufficient rate so as to provide adequate shrink fit stress to keep the disk 16 attached to the shaft 20. In accordance with the invention, a coating matrix 44 for enhancing thermal conductivity is applied at the interface 24 to increase heat transfer from the disk 16. The coating matrix 44 includes a graphene layer. Graphene has a thermal conductivity coefficient of up to 5300 W·m⁻¹·K⁻¹, which is approximately 300 times greater than that of stainless steel and approximately 20 times greater than copper. The coating matrix 44 may be applied to either the bore 22 or the shaft 20, or both the bore 22 and shaft 20, in order to maintain the shrink fit arrangement during fast start-up.

Referring to FIG. 5, a cross sectional view of the disk 16 and shaft 20 is shown which depicts additional applications of the coating matrix 44. The coating matrix 44 may also be formed on first 25 and second 27 side surfaces, the outer diameter 21 and the grooves 23 of the disk 16 in order to improve the transfer of heat in a radial direction along the disk 16. In particular, heat is transferred in a radial direction from hot areas of the disk 16 located along an outer radius 29, where heat may cause damage, to an area near an inner radius 31 where the disk 16 is susceptible to fracture due to relatively low temperatures that occur during start-up of a turbine. The coating matrix 44 may also be applied to disk 16 and shaft 20 assemblies which do not use a shrink fit arrangement to attach the disk 16 to the shaft 20 and to disk 16 and shaft 20 assemblies used in steam turbines and other types of engines.

Referring to FIG. 6, a cross sectional view of the disk 16 is shown without the shaft 20. The previously described bore 22 includes a bore surface 48 that contacts the shaft 20. The disk 16 may be fabricated from forged steel or a nickel (Ni) based superalloy. The bore surface 48 includes a receiving section 50 for receiving the coating matrix 44. The coating matrix 44 has a thermal conductivity which is greater than either the disk 16 or the shaft 20. In one embodiment, a first coating matrix 52 includes a graphene layer 54 that is directly attached to the receiving section 50 as shown in FIG. 7. Referring to FIGS. 7 and 6, the graphene layer 54 may be attached to the receiving section 50 by first melting the receiving section 50. In particular, the receiving section 50 is melted to a depth that is approximately equivalent to a thickness of the graphene layer 54 that is to be deposited on the receiving section 50. For example, the receiving section 50 may be melted to a depth of approximately 10-1000 micrometers. The graphene layer 54 is then deposited onto the receiving section 50 to attach the graphene layer 54 to the receiving section 50. A heat generating system 56 may be used to melt the receiving section 50. The heat generating system 56 may be a known pulsed laser system, which enables control of the amount of energy that is transmitted to the receiving section 50. The pulsed laser system melts the surface locally in a defined region on the bore surface 48, such as the receiving section 50. Further, laser parameters such as laser spot size, intensity and pulse length are selected so as to provide a desired melting of the receiving section 50. Alternatively, a known robot welding system may be used instead of the pulsed laser system to melt the receiving section 50. In another embodiment, the graphene layer 54 is first deposited on the receiving section 50. The receiving section 50 is then melted after the graphene layer 54 is deposited. The graphene layer 54 has a higher melting point relative to the material used to fabricated the disk 16 and thus is not melted by the heat generating system 56 when this approach is used.

In another embodiment, a second coating matrix 58 is formed on the receiving section 50 as shown in FIG. 8 by a known thermal spraying system 60 (see FIG. 6). The second coating matrix 58 includes a mixture of copper as will be described, or other material which is compatible with the disk 16, and a graphene layer 54 or multiple graphene layers 54. Alternatively, the second coating matrix 58 includes graphene and an anti-corrosion coating. In yet another embodiment, a graphene layer 54 is first deposited on the receiving section 50. Subsequently, a third coating matrix 62 that includes copper or an anti-corrosion coating is sprayed on the graphene layer 54. It is noted that multiple layers of the coating matrixes 52, 58, 62 may be sprayed on the receiving section 50.

Referring to FIGS. 9-10, alternate embodiments of the coating matrix 44 are shown. The coating matrix 44 may include a single graphene layer 54 or multiple graphene layers 54. In addition, a single copper layer 64 or multiple copper layers 64 may be added to the graphene layers 54 in order to further enhance thermal conductivity. For example, a copper layer 64 may be formed between graphene layers 54 to form an alternating graphene 54-copper 64 pattern as shown in FIG. 9. Use of an alternating graphene 54-copper 64 pattern provides enhanced adhesion of the graphene layers 54 to a surface such as the bore surface 48. Further, multiple graphene layers 54 may be located adjacent to each other to increase heat transfer as needed. For example, a copper layer 64 may be formed between a graphene layer 54 and adjacent graphene layers 54 as shown in FIG. 10. It is understood that the coating matrix 44 may include other graphene 54-copper 64 layer combinations. Alternatively, sufficient adhesion may be achieved by using sufficiently thin graphene layers 54, thus obviating the need for copper layers 64.

The coating matrixes 52, 58, 62 can also be applied to other areas or sections of a turbine as needed. For example, the coating matrixes 52, 58, 62 may be applied to non-shrunk-on structures or components in order to enhance the removal of heat, or alternatively to transport heat to areas where heat is needed such as in Hirth-serration couplings used in a turbine. Additionally, the coating matrixes 52, 58, 62 may be used for components or structures located in a hot gas path of the turbine, such as a turbine blade or vane, to guide or steer heat distribution. This enables the transfer of heat from critical hot areas of such components, which are prone to creep or other forms of service related degradation, to less critical areas.

The current invention may be applied to already existing turbines with minimal downtime and enables potential service upgrades without a substantial redesign of existing turbines. Further, the graphene provides corrosion and oxidation protection for a surface of a disk 16 or other components located in a hot gas path. This is an advantage for turbines exposed to difficult environmental conditions such as salty air found near an ocean or air that contains granular material such as sand particles found in desert environments.

While particular embodiments of the present invention have been illustrated and described, it would be obvious to those skilled in the art that various other changes and modifications can be made without departing from the spirit and scope of the invention. It is therefore intended to cover in the appended claims all such changes and modifications that are within the scope of this invention. 

What is claimed is:
 1. A method for forming a coating matrix on a component used in a shrink fit arrangement wherein the coating matrix has a higher thermal conductivity than the component, comprising: providing a receiving section on the component; heating the receiving section to melt the receiving section; and depositing at least one coating matrix layer on the receiving section.
 2. The method according to claim 1, wherein the receiving section is melted to a depth approximately equivalent to a thickness of the coating matrix layer.
 3. The method according to claim 2, wherein the depth is approximately 10-1000 micrometers.
 4. The method according to claim 1, wherein the coating matrix includes graphene.
 5. The method according to claim 1, wherein the coating matrix includes graphene and copper.
 6. The method according to claim 1, wherein the coating matrix includes a graphene layer located between copper layers.
 7. The method according to claim 6, wherein the coating matrix includes graphene layers which are adjacent each other.
 8. The method according to claim 1, wherein the coating matrix is applied to a shrink fit interface between shrink fit components.
 9. The method according to claim 1, wherein the receiving section is located on a bore surface of a turbine disk.
 10. The method according to claim 1, wherein the receiving section is located on a surface of a turbine shaft.
 11. A method for forming a coating matrix on a component used in a shrink fit arrangement wherein the coating matrix has a higher thermal conductivity than the component, comprising: providing a receiving section on the component; thermal spraying a graphene layer on the receiving section; and thermal spraying a copper layer on the graphene layer thereby forming the coating matrix.
 12. The method according to claim 11 further including thermal spraying an anti-corrosion coating on the receiving section.
 13. The method according to claim 11, wherein a graphene layer is formed between copper layers.
 14. The method according to claim 13, wherein at least two graphene layers are formed adjacent each other.
 15. The method according to claim 11, wherein the coating matrix is applied to a shrink fit interface between shrink fit components.
 16. The method according to claim 11, wherein the receiving section is located on a bore surface of a turbine disk.
 17. The method according to claim 11, wherein the receiving section is located on a surface of a turbine shaft.
 18. A shaft and disk assembly for use in a turbine, comprising: a disk having a central bore; a shaft received by the central bore, wherein the shaft and central bore are fastened by a shrink fit arrangement; and a coating matrix located between the central bore and the shaft, wherein the coating matrix has a higher thermal conductivity than the disk.
 19. The shaft and disk assembly according to claim 18, wherein the coating matrix includes graphene.
 20. The shaft and disk assembly according to claim 18, wherein the coating matrix is formed on a receiving section of a surface of the central bore.
 21. A shaft and disk assembly for use in a turbine, comprising: a disk having a central bore and an outer diameter which includes a plurality of grooves, wherein each groove is adapted to receive an associated blade; a shaft affixed to the central bore; and a coating matrix applied to first and second sides of the disk, the outer diameter, the grooves and between the central bore and the shaft, wherein the coating matrix has a higher thermal conductivity than the disk.
 22. The shaft and disk assembly according to claim 21, wherein the coating matrix includes graphene. 